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RESEARCH MEMORANDU M
DESIGN OF COMBUSTOR FOR LONG-RANGE RAM- J ET ENGINE AN D PERFORMANCE OF RECTANGULAR ANALOG By Warren D. Rayle and Richard G. Koch Lewis Fli ght Propulsion Laboratory Cleve land, Ohio This material conlalns lnformallon affecting 1 Defense of the Unlted Slates wlthln the meaning of the espionage laws. TIlle 18, U.S.C .• Sees. 7Q the transmtsslon or revelation of which In any manner to an unauthorized person 1s prohibited w.
NATIONAL AD SORY COMMITTEE
FOR AERONAUTICS
WASHINGTON January 29, 1954
r
NACA RM E5 3K13 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Iill3EARCH MEMORANDUM DESIGN OF COMBUSTOR FOR LONG-RANGE RAM-JET ENGINE AND PERFORMANCE OF RECTANGULAR ANALOG By Warren D. Rayle and Richard G. Koch SUMMARY This report describes the design of a piloted combustor intended for a ram-jet engine of long flight range. The unit comprises a large annular basket of V-type cross section, the inner surface of which is slotted and bent into small V-gutters. At the trailing edge of the basket, eight V-gutters are used to propagate the flame into the main stream.
A rectangular analog of this combustor was tested at air-flow con- ditions corresponding to tho ' se that might be obtained during cruise. At these conditions, combustion efficiencies of as much as 90 percent were calculated for the combustor at the design equivalence ratio of 0.52.
The performance of the unit was relatively insensitive to mounting and flow variables; the greatest effect on efficiency was that of the manner and location of the fuel injection.
A full-scale version of this combustor has been designed for a 48-inch-diameter engine.
INTRODUCTION The following material describes the design and testing of a ram- jet combustor intended to operate at a specific set of flow conditions corresponding to those that might be obtained in a long-range ram-jet engine. The work was performed at the NACA Lewis laboratory as part of a continuing program of combustor design and evaluation.
The design and testing were based on an engine of 48-inch diameter and 75-inch length, with the shape shown in figure l(a). The altitude and flight Mach number were assumed to be those required to supply the following combustor-inlet conditions: (1) a total pressure of 10 inches of mercury absolute, (2) a temperature of 530 F, and (3) a Mach number of 0.15. The goal of the program was to attain maximum efficiency at NACA RM E53K13 these conditions for an equivalence ratio of 0.52 wit h out introducing an internal total-pressure loss of more than 3 velocity heads. Obviously, if the combustor is to operate under realistic conditions, the perform- ance must remain high for small deviations from the design conditions.
Combustor development and testing in an engine as large as 48 inches in diameter would obviously be cumbersome and inefficient. A 10- by 24-inch rectangular analog was therefore fabricated. This test combustor was designed to correspond in length to the full-size engine and to induce similar air velocities at all stations.
The simulation extended from the throat of the inlet diffuser to the exhaust choke, thus isolating the unit acoustically. The side view of the analog is presented in figure l(b). A rectangular cross section exists at all stations, with the height of the rectangle equal to the radius of the engine.
A kerosene-type fuel was used throughout. The properties of this fuel, MIL-F-5616 grade JP-l, are given in table I.
Any flame-holding device functions by providing a continuous source of ignition for gases flowing at velocities greater than the normal laminar flame speed (about 1 to 2 ft/sec). The baffle-type flame holder accomplishes this by a trailing vortex which recirculates the hot com- bustion gases. A can-type combustor serves much the same purpose, although a separate fuel source is often provided and the vortex region is more sheltered from the velocity fluctuations occurring in the main stream. When a can-type combustor is used to provide ignition sources for a system of baffle-type flame holders, the can is frequently referred to as a "pilot."
In the past, can-type flame holders have been used in engines in which stable performance and high efficiency were desired at low pres- sures and in which a high drag was permissible (ref. 1). Baffle-type flame holders have been used when the drag must be kept low and when the added stability of the can was not needed (ref. 2). For the conditions of these tests, neither type alone seemed adequatej therefore, a combina- tion was explored. A protected region, or pilot, with its own fuel supply was contrived to provide a continuous source of ignition for a V-gutter flame holder, which, in turn, spread the flame throughout the main stream. The joining of the main stream to the pilot flame was made gradual, so that the continuity of the flame might not be easily broken.
Since the desired fuel-air ratio was only about half of stoichiometric, an annular pilot was designed to serve also as a flow divider. This design permitted the main combustion to occur in stoichiometric mixtures in the central portion of the duct, as in the combustor of reference 3.
Figure 2 shows this combustor mounted in the rectangular duct and fig- ure 3 shows the corresponding full-scale unit.
__ I
NACA RM E53K13 Inasmuch as the combustion-chamber length was only 75 inches, the pilot burner was mounted in the subsonic diffuser so that full advantage might be taken of the available reaction zone. This method 01' mounting should assist the performance of , the diffuser by reducing the tendency toward flow separation and therefore should not too greatly increase the drag.
The performance of this unit was evaluated at design conditions.
Data were also taken to ascertain the effect of the individual variables - pressure, temperature, flow velocity, mounting angle, pilot fuel flow, location of main fuel injection, and vitiation of the inlet air.
APPARATUS Test Facility An outline of the combustor and the associated ducting used in these tests is presented in figure l ( b) . Air at 40 pounds per square inch gage was supplied by the laboratory facilities; the flow rate was controlled by remotely operated butterfly valves . The electric preheaters were not capable of heating the mass of air (5 lb/sec at simulated cruise condi- tions) to the desired temperature and were supplemented by a single J35 turbojet combustor through which part of the air was passed, burning and subsequently mixing with the main stream. In order to increase the effi- ciency of the preheater, the pressure therein was maintained at an ele- vated value by means of a fixed - area orifice downstream. After passing through a plenum chamber, the air was introduced into the engine analog by means of a two-dimensional duct which increased in cross section from 5 by 12 inches at the inlet to a maximum of 10 by 24 inches.
A 6-inch-square window in the plenum chamber afforded a view of the combustion process.
Pilot Combustor The pilot fuel was introduced within the basket; the main fuel was sprayed into the region below. An oxygen-hydrogen flame was used as an igni ter, and the ensuing combustion occurred wi thi O n and downstream of the pilot. The main combustion region was located in the 10- by 24-inch duct fabricated from 1/2-inch steel plate. No forced cooling was em- ployed. The exhaust choke was constructed from a series of 3/4-inch tubes through which quench water was sprayed to halt the reaction. A grid of tapered Inconel members moved into the spaces between the quench tubes, thus providing a variable-area exhaust nozzle. This nozzle did not correspond to that of the engine but was necessary in order to permit the testing to be conducted at constant inlet pressure, temperature, and 4 NACA RM E53K13 Mach number while the fuel-air ratio was varied. The mixture of com- bustion products, air, and water vapor then passed through an array of thermocouples and into the exhaust mains.
The first pilot configuration tested is shown in figure 2. In cross section its shape was that of an asymmetric V. The outer surface was perforated by four rows of 1!2-inch holes spaced 2 inches apart.
Downstream of these holes the metal sheet was continued in order to act as a flow divider and to further protect the pilot zone. The total length of the upper surface was 24 inches. The lower surface of the basket was cut longitudinally, the cuts being spread at the downstream edge by folding the metal into a V-shape. The resulting openings per- mitted a gradual mixing of the main-stream gases with the pilot flame and also provided additional air entry for pilot combustion. From the center of the lower trailing edge of the pilot basket was appended a V-gutter, a flame seat for the main combustion. This basic configura- tion was used with slight modifications throughout the series of tests.
Instrumentation The mass flow of air was calculated from the pressure drop across a variable-area orifice. Fuel flow was determined from rotameter read- ings, the preheater fuel supply being measured independently from that of the experimental combustor. Two arrangements were used for con- trolling the pilot fuel flow. In the earlier tests, the pilot- and main-fuel systems were manifolded so that the flows remained roughly proportionate. Later, a separate control was used for the pilot-fuel system. In this case, the sum of the two flows was measured, and the pilot flow was estimated from the injection pressure. The quench-water flow was also determined from rotameter readings. All rotameters were calibrated after installation. The quench spray, an air-atomizing type, required a small flow of high-pressure air; this flow wa s meas- ured by use of a small fixed orifice.
Pressures at various points along the system were transmitted to mercury manometers whose readings were photographically recorded. Total and static pressures were obtained near the throat of the diffuser (sta- tion 1) and just upstream of the combustor (station 2). The combustor- inlet pressure was considered to be the total pressure at station 2.
Static pressures were also measured at the top and the side of the burner section and at three points on the side of the duct just ahead of the exhaust choke. Pressure taps downstream permitted measurement of the pressure drop across the choke.
Temperatures were measured by means of thermocouples strategically located along the system. Chromel-alumel thermocouples, connected to a recording instrument, were located as follows: (1) a single thermocouple NACA RM E53K13 at station 0, (2) five thermocouples spot -w elded to the combustor wall, and (3) 16 thermocouples arrayed at centers of equal areas in the duct at station 4. Iron-constantan thermocouples were employed to measure the temperatures of the air entering the preheater and of the quench water.
PROCEDURE Operation The running procedure finally evolved was as follows: First, the required mass flow of air was established. Then, the preheater was started and its fuel flow adjusted until the plenum - chamber temperature reached equilibrium at about 530 F. The pilot fuel was admitted and was ignited by an oxygen-hydrogen torch, followed by the main fuel flow and combustion. The combustor pressure was set to the required value of 10 inches of mercury absolute, and the quench spray was adjusted to give 0 0 a mean exhaust temperature between 400 and 600 F . Data were then taken.
The wall temperatures were recorded twice over a measured time interval (about 30 sec) in order to establish the rate of heat absorption of the walls. Another fuel flow was set and the process repeated. In general, a limit of four or five consecutive data points was enforced by over- heating of the combustor wall· When the maximum wall temperature 0 0 reached 1000 to 1100 F, the fuel fl ow was stopped and the combustor was permitted to cool.
Most of the data were taken at the simulated cruise condition and translated into curves of efficiency against equivalence ratio. Data on combustion limits were also sought at the standard flow conditions.
In general, the fuel system would not provide sufficient fuel to attain rich blow-out, the limitation being in the capacity of the main fuel nozzles. Lean limits were virtually nonexistent since the pilot basket retained flame even in the absence of the main fuel. The lean limit of the pilot itself was likewise indefinite; frequently, flame remained visible in the basket for as long as 10 to 20 seconds after the supply valve was closed.
Calculation Methods The thermal efficiency of the combustion process was deduced from a heat balance. The total heat content of the exhaust gases - air, superheated water vapor, and combustion products - was computed. Sim- ilarly, the heat content of the ingredients was summed. The ingredients included air, preheater combustion products, water, fuel, and quench air. The heat liberated in the unit was then the difference of these values plus the empirical ly determined loss from the combustor wall . The NACA RM E53K13 combustion efficiencies cited herein are then the ratio of heat liber- ated (including estimated loss) to the theoretical heating value of the fuel.
The equivalence ratios used on the plots were computed from the total air flow, fuel flow, preheater fuel flow, and preheater efficiency.
Thus, they include the vitiation resulting from preheater combustion.
In general, the operation of the preheater resulted in the pre consumption of about 5 percent of the available oxygen.
The heat loss of a unit such as this may be quite large. It com- prises three components - convection, radiation, and capacitance. The first two components are functions of the wall temperature; the storage term is a function of the rate of change of the wall temperature. In order to determine the relative magnitude of these terms, a set of data was taken with only the preheater used. With a constant fuel flow to the preheater, the heat loss of the gases passing through the combustor was measured periodically, as was also the wall temperature. From these data, an empirical heat-loss equation was calculated. This method of estimating losses, though not precise, was deemed sufficient for the purpose. Losses computed ranged from 5 to 20 percent, being smaller percentagewise at the higher equivalence ratios. The final effici ency figures with the loss included were reproducible to within about 2 percent.
The total-pressure loss was calculated from the total pressure at the inlet, the static pressure at the exit, and the total heat liberated.
This calculation involved the assumption that the temperature profile of the gas entering the exhaust choke was flat, since the temperature was computed from the mass flow of air and the heat liberated. The results indicate the magnitude of the pressure drop but are not sufficiently refined to permit comparisons of similar configurations.
RESULTS At an inlet velocity and temperature corresponding to cruise condi - tions, pilot combustion persisted until the pressure was reduced to about 6 . 5 inches of mercury absolute. Exhaustive tests of the stability limits were omitted, since the data quoted seemed to indicate sufficient sta- bility to justify transferring attention to the performance of the com- bined pilot and main combustor. Subsequently, the trailing V-gutter was added and the main fuel was injected.
Early tests of the original combustor yielded efficiency maximums at very low equivalence ratiOS, as shown in figure 4. In order to shift this peak to richer regions, the air flow through the pilot was increased by enlarging the second row of holes from 1/2- to 7/8-inch diameter. The RM E53K13 NACA unit of the efficiency 4. The on figure also shown data are resulting to about 79 percent from about dropped ratio equivalence at design a single through introduced fuel was the main both tests For percent.
and downstream, directed per hour, 60 gallons rated at nozzle swirl trail- of the lower upstream 5 inches below and l~ inches about located flow con- off-design at were obtained The data the pilot.
edge of ing inlet total-pressure of the failure of vibration as a result ditions the static pressure, and inlet flow, temperature, the mass From tube.
were calculated inlet Mach number and total pressure values of correct 4.
on figure as shown to be the pilot through injected of fuel the amount of varying The effect from resulted in efficiency change No significant figure 5.
shown in is equiv- the term "pilot 0.103; ratio to equivalence the pilot increasing fuel theoretically fuel to the of pilot the ratio describes ratio" alence This combustor.
the total to air supplied all the to burn necessary pilot, of the efficiency the constant to be attributed may insensitivity and both before presented data are 6. The on figure is shown which that in be noted It may included.
was correction the heat-loss after a high by was accompanied liberation of heat a low rate instance this ratio of at an equivalence data point The of heat loss.
percentage the series.
point of the first curve is fallon the does not which 0.085 of uneven distribution to the attributed may be this point error at The of an calculation in the resulted wall, which combustor in the heat heat loss.
high unduly r~ efficiency for maximum ratio equivalence over-all Because the the prob- to reduce In order were made.
changes low, further too mained fuel- the rich regions, in locally occurring combustion of ability gallons at 40 nozzles rated swirl Two was modified.
system injection trailing basket's of the inches upstream about 16 located hour and per nozzle the Single to replace were used equal areas of edge at centers This change upstream.
5 inches located hour and gallons per at 60 rated to 85 76 percent from about condition at design efficiency the raised at an equivalence 89 percent of efficiency gave a maximum and percent of is the effect shown 7. Also on figure as is shown of 0.42, ratio angle between is, the pilot, that of the of attack the angle increasing An of the combustor.
line and the base the pilot line of the center 0 the performance.
affect not greatly to 8 did from 6 increase 8 shows Figure next varied.
was fuel injection of the direction The bar.
the fuel-injection by rotating merely obtained effect the drastic and upstream representing are those interest of greatest curves The two peak to shift the found was injection Upstream injection.
downstream represent- The curve ratio.
equivalence desired about the to efficiency it indicates interest; practical of little is injection ing cross-stream fuel spray the from directing resulting efficiency in rich the decline by be explained of course, may, This result the pilot.
toward upward, mixtures.
rich in locally results arrangement that this assuming NACA RM E53K13 The effect of combustor-inlet pressure is shown in figure 9 (the fuel being injected upstream). Increasing the pressure from 10 to 12 inches of mercury raised the efficiency at the design condition from about 91 to 95 percent; lowering the pressure to 8 inches dropped the efficiency to about 73 percent and shifted the efficiency peak to a richer region. Brief tests of the final design at an inlet pressure of 17.5 inches of mercury showed no resonant instability, no overheating of the basket, and efficiency on the order of 95 percent.
The utility of the flow divider (the extension of the outer surface of the basket) was next examined. Stepwise removal of this surface indicated that an excess of metal was originally present. Figure 10 presents efficiency curves for the original basket compared with those for the same basket with 6 and 12 inches of the outer surface removed.
The first modification is seen to perform as did the original. The removal of 12 inches of the metal reduced the stability of the unit so that blow-out occurred at an e~uivalence ratio of about 0.47.
These tests were all conducted with combustion preheat; that is, a small amount of the available oxygen had been already consumed. In order to determine whether this factor would seriously affect the re- sults, the data represented by figure 11 were obtained. It should be noted that these data were taken with an early configuration (single nozzle for the main fuel) and not with the final design. The trends should still apply . In order to vary the amount of vitiation with o ut changing the inlet temperature, the same conditions were obtained with and without electric preheat by changing the amount of combustion pre- heat. The chief effect of the vitiation seems to be to increase the slope of the curve . Surprisingly enough, at the lean condition the per- formance was actually better with vitiated air. This improvement was slight and may reflect data scatter. The effect at the design e~uivalence ratio was slight but adverse. It may be tentatively concluded that the vitiation normally present during these tests would cause the results to be, if anything, conservative.
The total-pressure drop through the duct with combustion was com- puted for several instances to be about 6 velocity heads. This drop seemed excessive and was checked by removing the combustor and acces- sories. The empty duct alone was found to have a drag of about 3 veloc- ity heads. Therefore, it was assumed that the drag due to the combustor itself might approach 3 velocity headsj this matter may be more accu- rately checked with the full-scale model.
CONCLUDING REMARKS From the results shown in the preceding section, it may be con- cluded that the combustor developed very nearly satisfies the original RM E53K13 NACA at obtained percent was about 90 of peak efficiency The requirements.
was near The drag of 0.52.
ratio equivalence operating the required con- operation efficient permitted stability and the heads, 3 velocity cannot ensure The data conditions.
from the standard removed siderably pressures occur at the will not or both, burnout, or that instability and for design a basis do provide but they altitude, at low encountered testing.
full-scale insensi- is configuration particular that this indicate The results of fuel injec- manner on the dependent is strongly but tive to mounting the two-dimensional of the conversion facilitates greatly tion. This areas, as the open inasmuch basket, annulus a full-scale design to those in with identical be maintained cannot all and angles lengths, a Such adjusted.
is mOre readily injection fuel unit, whereas the test was designed.
combustor full-scale Laboratory Flight Propulsion Lewis for Aeronautics Committee Advisory National 17, OhiO, November Cleveland, REFERENCES Com- of Ramjet Development H. W.: McFarland, C. L., and 1. Dailey, Calif., Univ. Southern Aero. Lab., No. 13-5, Rep.
USCAL ponents.
U.S.
Sept., 1951, Aug. & June, July, Rep. for (Prog.
8, 1951.
Oct.
.
NOas 51-116-c.)
Contract Bur. Aero.
Navy, Altitude- G., Jr.: Henzel, J.
B., and T.
W. L., Shillito, 2. Jones, Ram-Jet a 28-Inch of of Performance Investigation Test-Chamber of Four Performance and Operational - Combustion I Engine.
1950.
RM E50F16, NACA Configurations.
Combustion-Chamber of Fuel-Air E. E.: Effect Dangle, D. W., and Bahr, A. J., 3. Cervenka, Performance on Combustion in Combustion Zone Concentration Ratio 1953.
RM E53B19, NACA Engine.
16-Inch Ram-Jet of a NACA RM E53K13 TABLE I. - ANALYSIS OF MIL-F-5616 GRADE JP-l FUEL A.S.T.M. distillation Initial boiling point, ~ Percentage evaporated 60 361 70 370 80 384 90 406 95 424 Final boiling point 458 Residue, percent 1. 0
Loss, percent o
Aromatics, percent 14 Specific gravity 0.796 Hydrogen-carbon ratio 0.163 Net heat of combustion, 18,59 5 Btu/lb ~ &; ~ t>:.1 U1 ~ I--' (fl I--' I--' exhaust : t water exhaus engine.
Laboratory quench > ducting.
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full-scale I I associated 48-inch ~
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